The present invention relates to a gas turbine combustor and to a structure for reducing the disturbances in an air flow in the combustor so that the combustion instability may be reduced.
FIG. 13 is a general sectional view of a gas turbine. In FIG. 13, numeral 1 designates a compressor for compressing air for combustion and for cooling a rotor and blades. Numeral 2 designates a turbine casing, and numeral 3 designates a number of combustors arranged in the turbine casing 2 around the rotor. There are for example sixteen combustors, each of which is constructed to include a combustion cylinder 3a, a cylinder 3b and a transition cylinder 3c. Numeral 100 designates a gas path of the gas turbine, which is constructed to include multistage moving blades 101 and stationary blades 102. The moving blades are fixed on the rotor and the stationary blades are fixed on the side of the turbine casing 2. The hot combustion gas, jetting from the combustor transition cylinder 3c, flows in the gas path 100 to rotate the rotor.
FIG. 14 is a detailed view of portion 3a in FIG. 13 and shows the internal structure of the combustor 3. In FIG. 14, numeral 4 designates an inlet passage of the combustor, and numeral 5 designates a main passage or a passage around main nozzles 7. A plurality of, e.g., eight main nozzles 7 are arranged in a circle. Numeral 6 designates a main swirler which is disposed in the passage 5 of the main nozzles 7 for swirling the fluid flowing in the main passage 5 toward the leading end. Numeral 8 designates one pilot nozzle, which is disposed at the center and which is provided therearound with a pilot swirler 9, as in the main nozzles 7. Numeral 10 designates a combustion cylinder.
In the gas turbine combustor thus far described, the air, as compressed by the compressor 1, flows, as indicated by 110, from the compressor outlet into the turbine casing 2 and further flows around the inner cylinder of the combustor into the combustor inlet passage 4, as indicated by 110a. After this, the air turns around the plurality of main nozzles 7, as indicated by 110b, and flows into the main passage 5 around the main nozzles 7, as indicated by 110c. On the other hand, the air also flows around the pilot nozzle 8, as indicated by 110d, and is swirled by the main swirler 6 and the pilot swirler 9 until it flows to the individual nozzle leading end portions, as indicated by 110e, for combustion.
FIG. 15 is a diagram showing the flow states of the air having flowed into the combustor of the prior art. The air 110a from the compressor flows, as indicated by 110b, around the main nozzles 7. Around the outer sides of the main nozzles 7, however, vortexes 120 are generated by the separation of the flow. When the air flows in from the root portion around the pilot nozzle 8, on the other hand, there are generated vortexes 121, vortexes 122 flowing to the leading end of the pilot nozzle 8, and disturbances 123 in the flow around the outlet of the inner wall of the combustor.
In this gas turbine, NOx is emitted more as the load becomes heavier, but this emission has to be suppressed. As the load is increased, the combustion air has to be increased accordingly. As described with reference to FIG. 15, the air vortexes 120, 121, 122 and 123 in the combustor are more intensified, increasing the tendency to combustion instability. In order to suppress the emissions of NOx, the aforementioned combustion instability is reduced at present by adjusting the pilot fuel ratio and the bypass valve opening. With the prevailing structure, however, the running conditions are restricted by the combustion instability.
In the gas turbine combustor of the prior art, as has been described hereinbefore, drifts, vortexes and flow disturbances are caused in the air flowing in the combustor, causing the combustion instability. As the load is increased, increasing the flow rate of air to combustion, so that the drifts, vortexes and flow disturbances have serious influences, the concentration of the fuel becomes heterogeneous with respect to time and space, thereby making the combustion unstable. At present, in order to suppress this combustion instability, the pilot combustion ratio and the bypass valve opening are adjusted, but in vain for sufficient combustion stability. In the worst case, therefore, the combustor is damaged and the gas turbine running range is restricted.
Therefore, the present invention has been conceived to provide a gas turbine combustor which reduces combustion instability by guiding the air to flow smoothly into the combustor and by straightening the flow to eliminate flow disturbances and concentration changes of the fuel.
In order to solve the foregoing problems, the present invention contemplates the following.
(1) A gas turbine combustor comprises a cylinder supported at its circumference by a plurality of struts fixed on one end in a combustor housing portion of a turbine casing. A pilot nozzle is arranged at the center of the cylinder. A plurality of main nozzles are arranged around the pilot nozzle. A flow ring has a ring shape so as to cover the upstream end of the cylinder, a semicircular sectional shape (including an elliptical shape) and so maintains a predetermined gap. A porous plate downstream of the flow ring closes a space which formed in the cylinder between the pilot nozzle and the main nozzles.
(2) A gas turbine combustor as set forth in (1), can have the flow ring sectionally shaped as an extended semicircular shape by extending the two ends of a semicircle. The porous plate is fixed at its circumference on the circumferential side face of the extended semicircular shape.
(3) A gas turbine combustor as set forth in (1), can have the flow ring include semicircular curves arranged in multiple stages while maintaining a predetermined gap.
(4) A gas turbine combustor as set forth in (1), can have a guide portion disposed around the inlet portion of the combustor housing portion of the turbine casing with a smoothly curved face for covering the whole circumference wall face of the inlet portion.
(5) A gas turbine combustor as set forth in (1), can also have a funnel shaped flow guide having a smoothly curved sectional shape along the curved face of the flow ring and arranged upstream of the flow ring while maintaining a predetermined gap from the flow ring. The flow guide is fixed at its larger diameter portion on the inner wall of the combustor housing portion of the turbine casing and at its smaller diameter portion around the pilot nozzle. The porous plate is arranged downstream of a support for supporting the pilot nozzle and the main nozzles.
(6) A gas turbine combustor according to the present invention may also comprise a cylinder supported at its circumference by a plurality of struts fixed on one end in a combustor housing portion of a turbine casing. A pilot nozzle is arranged at the center of the cylinder. A plurality of main nozzles are arranged around the pilot nozzle. A flow ring having a ring shape covers the upstream end of the cylinder, has a semicircular sectional shape and maintains a predetermined gap. Flow rings individually having semicircular sectional shapes are arranged in multiple stages upstream of the flow ring in the axial direction while maintaining a predetermined gap. A cylindrical porous plate covers the entire circumference of the inlet portion on the outer side of all of the flow ring.
(7) Another gas turbine combustor may comprise a pilot nozzle arranged at the center of a cylinder and a plurality of main nozzles arranged around the pilot nozzle. Spaces between the circumference of the pilot nozzle and the inner circumferences of the individual main nozzles confronting each other are filled with a filler in the axial direction downstream from the upstream end so as to extend near the circumferential portion of the leading end of the cylinder, thereby forming fairings. The passage between the adjoining fairings is made wider on the downstream side than on the upstream side.
(8) Another gas turbine may comprise a compressor and a combustor, the combustor comprising a cylinder supported at its circumference by a plurality of struts fixed on one end in a combustor housing portion of a turbine casing. A pilot nozzle is arranged at the center of the cylinder and a plurality of main nozzles are arranged around the pilot nozzle. A flow guide is disposed around the entire circumference of the outlet of the compressor, having a smoothly curved face for guiding the discharged air to flow toward the combustor on the outer side. The combustor comprises a flow ring having a ring shape so as to cover the upstream end of the cylinder with a semicircular sectional shape and so as to maintain a predetermined gap. A porous plate is arranged downstream of the flow ring for closing a space which is formed in the cylinder between the pilot nozzle and the main nozzles. A guide portion has a smooth curved face and is disposed around the inlet portion of the combustor housing portion of the turbine casing for covering the entire circumference wall face of the inlet portion.
In the invention (1), the air to flow in the combustor flows at first smoothly along the curved face of the flow ring in the cylinder and then passes through the numerous pores of the porous plate so that it is straightened into a homogenous flow. With neither separation vortexes nor flow disturbances, unlike the prior art, the air flows along the pilot nozzle and the main nozzles to the leading end portion so that combustion instability, as might otherwise be caused by the concentration difference of the fuel, can be reduced.
In the invention (2), the flow ring is formed into an extended semicircular shape, and the porous plate can be fixed at its periphery on the extended semicircular side face so that manufacture can be facilitated. In the invention (3), on the other hand, the flow rings are arranged in multiple stages so that the air is homogeneously guided to flow into the cylinder of the combustor through the multistage circumferential gaps to thereby better promote the effects of invention (1).
In the invention (4), the inlet portion of the combustor housing portion for the air is constructed of the wall faces having corners for protruding into the housing portion. The air to flow into the combustor is disturbed and is guided in a turbulent state into the flow guide of the leading end portion of the combustor. However, the guide portion is provided so that the wall face of the inlet portion may form a smoothly curved face. With this guide portion, the air inflow can be prevented from being disturbed, reducing the combustion instability as with invention (1).
In the invention (5), the air inflow is smoothly turned at the upstream end of the combustor by the funnel-shaped flow guide and is guided into the cylinder by the flow ring. Moreover, the porous plate is disposed downstream of the support for supporting the pilot nozzle and the main nozzles. Even if the flow is more or less disturbed by the support therefore, these disturbances are straightened by the porous plate so that the air flow is homogenized and introduced into the nozzle leading end portion to thereby better ensure the reduction of combustion instability.
In the invention (6), the flow rings are arranged in multiple stages, and the cylindrical porous plate is arranged in front of the air inlet portion around those flow rings. Therefore, the air to flow into the combustor is straightened into cylindrical homogeneous flow by the porous plate. This homogeneous flow is then smoothly guided through the gap between the multistage flow rings into the cylinder of the combustor. In the invention (6), too, the disturbances of the air flow are reduced to reduce the combustion instability.
In the invention (7), in the space between the individual main nozzles and the pilot nozzle opposed to each other, the fairings are formed so that the air flows in the gaps between the adjoining fairings and further flows downstream. This air flow has a downward rising flow velocity. Therefore, the gap is enlarged from upstream to downstream so that the air flow through the fairings is homogenized by the shape. Thus, the air can flow downstream without any flow disturbance to thereby reduce combustion instability as might otherwise be caused by its disturbance.
In the invention (8), the flow guide is disposed at the compressor outlet for guiding the air flow from the compressor outlet to the combustor homogeneously around the combustor. In the combustor, the flow ring and the porous plate are disposed to eliminate the air disturbances in the combustor and to reduce combustion instability. Moreover, the air to flow in the combustor is guided to flow smoothly at the inlet portion of the combustor housing portion by the guide portion of the smooth curve. As a result, there can be realized a gas turbine which can reduce the pressure loss in the air flow and can reduce combustion instability.